U.S. patent number 10,450,074 [Application Number 15/205,648] was granted by the patent office on 2019-10-22 for aircraft pneumatic system.
This patent grant is currently assigned to ROLLS-ROYCE plc. The grantee listed for this patent is ROLLS-ROYCE plc. Invention is credited to Ivan Popovic, Rory D. Stieger.
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United States Patent |
10,450,074 |
Stieger , et al. |
October 22, 2019 |
Aircraft pneumatic system
Abstract
An aircraft pneumatic system includes a main gas turbine engine
including a main compressor, a recuperated auxiliary gas turbine
engine including a core compressor having an inlet in fluid
communication with a main compressor bleed of the main gas turbine
engine and an environmental control system inlet in fluid
communication with the main compressor bleed of the main gas
turbine engine. The environmental control system includes a
compressor driveable by power provided by the auxiliary gas turbine
engine.
Inventors: |
Stieger; Rory D. (Derby,
GB), Popovic; Ivan (Derby, GB) |
Applicant: |
Name |
City |
State |
Country |
Type |
ROLLS-ROYCE plc |
London |
N/A |
GB |
|
|
Assignee: |
ROLLS-ROYCE plc (London,
GB)
|
Family
ID: |
54200376 |
Appl.
No.: |
15/205,648 |
Filed: |
July 8, 2016 |
Prior Publication Data
|
|
|
|
Document
Identifier |
Publication Date |
|
US 20170106985 A1 |
Apr 20, 2017 |
|
Foreign Application Priority Data
|
|
|
|
|
Aug 7, 2015 [GB] |
|
|
1513952.0 |
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02C
9/18 (20130101); B64D 13/02 (20130101); F01D
15/10 (20130101); B64D 41/00 (20130101); H02K
7/1823 (20130101); B64D 13/06 (20130101); F02C
3/13 (20130101); F02C 7/277 (20130101); F02C
7/08 (20130101); F02C 6/08 (20130101); B64D
27/16 (20130101); Y02T 50/60 (20130101); B64D
2013/064 (20130101); F05D 2220/50 (20130101); F05D
2220/323 (20130101); Y02T 50/56 (20130101); B64D
2013/0618 (20130101); Y02T 50/50 (20130101); B64D
2013/0611 (20130101); Y02T 50/671 (20130101); F05D
2220/76 (20130101); Y02T 50/44 (20130101); B64D
2013/0648 (20130101); B64D 2013/0644 (20130101); Y02T
50/40 (20130101); B64D 2013/0603 (20130101) |
Current International
Class: |
B64D
13/02 (20060101); B64D 13/06 (20060101); F02C
6/08 (20060101); F02C 7/08 (20060101); F02C
7/277 (20060101); B64D 27/16 (20060101); B64D
41/00 (20060101); F01D 15/10 (20060101); F02C
3/13 (20060101); F02C 9/18 (20060101); H02K
7/18 (20060101) |
Field of
Search: |
;60/39.511,785,782 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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|
|
|
|
0888966 |
|
Jan 1999 |
|
EP |
|
1752377 |
|
Feb 2007 |
|
EP |
|
2204319 |
|
Jul 2010 |
|
EP |
|
2 871 349 |
|
May 2015 |
|
EP |
|
2881329 |
|
Jun 2015 |
|
EP |
|
2436708 |
|
Oct 2007 |
|
GB |
|
20130058849 |
|
Jun 2013 |
|
KR |
|
2013/121139 |
|
Aug 2013 |
|
WO |
|
Other References
Dec. 15, 2016 Search Report issued in European Application No. 16
17 8641. cited by applicant .
Jan. 27, 2016 Search Report issued in British Patent Application
No. 1513952.0. cited by applicant.
|
Primary Examiner: Jules; Frantz F
Assistant Examiner: Nouketcha; Lionel
Attorney, Agent or Firm: Oliff PLC
Claims
The invention claimed is:
1. An aircraft pneumatic system comprising: a main gas turbine
engine comprising a main compressor; a recuperated auxiliary gas
turbine engine comprising a core compressor, a combustor, and a
recuperator, the core compressor including an inlet in fluid
communication with a main compressor bleed of the main gas turbine
engine, and the recuperator including a heat exchanger that is
configured to recycle waste heat from an exhaust of the recuperated
auxiliary gas turbine engine to an auxiliary gas turbine engine
flow prior to the combustor; and an environmental control system
inlet in fluid communication with the main compressor bleed of the
main gas turbine engine, the environmental control system inlet
comprising a compressor drivable by power provided by the
recuperated auxiliary gas turbine engine.
2. The system in accordance with claim 1, wherein the inlet of the
core compressor is in direct fluid communication with the main
compressor bleed of the main gas turbine engine.
3. The system in accordance with claim 1, wherein the inlet of the
core compressor is in fluid communication with the main compressor
bleed of the main gas turbine engine via a cabin air exhaust.
4. The system according to claim 1, wherein the recuperated
auxiliary gas turbine engine comprises a variable area turbine.
5. The system according to claim 1, wherein the inlet of the core
compressor is in fluid communication with one or more of a lower
pressure main compressor customer bleed port, a higher pressure
main compressor customer bleed port, a fan outlet of the main gas
turbine engine, and a compressor handling bleed port.
6. The system according to claim 1, further comprising a valve
moveable between a first position, in which at least one of the
environmental control system inlet and the inlet of the core
compressor is in fluid communication with a fan outlet of the main
gas turbine engine, and a second position in which at least one of
the environmental control system inlet and the inlet of the core
compressor is in fluid communication with a compressor bleed of the
main compressor.
7. The system according to claim 1, wherein the recuperated
auxiliary gas turbine engine comprises a load driven by a turbine
of the recuperated auxiliary gas turbine engine, the load
comprising at least one of an electrical generator, a load
compressor, a main gas turbine engine fluid pump, a main gas
turbine engine shaft, and an air cycle machine compressor.
8. The system according to claim 7, wherein the load compressor is
configured to provide air to an inlet of the air cycle machine
compressor.
9. The system according to claim 8, wherein the system comprises a
selectable bypass arrangement configured to provide air to the air
cycle machine compressor directly from the main gas turbine engine
compressor when operated in a bypass mode, and from the load
compressor when in a non-bypass mode.
10. The system according to claim 7, wherein the electrical
generator is electrically connectable to an electric motor and the
air cycle machine compressor is drivable by the electric motor.
11. The system according to claim 10, wherein the air cycle machine
comprises a first compressor driven by the electric motor, and a
second compressor driven by the air cycle machine turbine.
12. The system according to claim 1, wherein the environmental
control system comprises an air cycle machine comprising, in fluid
flow sequence, a compressor, a heat exchanger and a turbine, the
compressor being drivable by the turbine.
13. The system according to claim 1, wherein the recuperated
auxiliary gas turbine engine is located either within a main gas
turbine engine nacelle or within an aircraft fuselage.
14. The system according to claim 13, wherein the recuperated
auxiliary gas turbine engine comprises a propelling exhaust nozzle,
which is located within the main gas turbine engine nacelle.
15. An aircraft pneumatic system comprising: a main gas turbine
engine comprising a main compressor; a recuperated auxiliary gas
turbine engine comprising a core compressor, a combustor, and a
recuperator, the core compressor including an inlet in fluid
communication with a main compressor bleed of the main gas turbine
engine, and the recuperator including a heat exchanger that is
configured to recycle waste heat from an exhaust of the recuperated
auxiliary gas turbine engine to an auxiliary gas turbine engine
flow prior to the combustor; an environmental control system inlet
in fluid communication with the main compressor bleed of the main
gas turbine engine, the environmental control system inlet
comprising a compressor drivable by power provided by the
recuperated auxiliary gas turbine engine; and a pre-cooler that
provides pre-cooled air to be combined with an air flow from the
recuperated auxiliary gas turbine engine.
16. An aircraft pneumatic system comprising: a main gas turbine
engine comprising a main compressor; a recuperated auxiliary gas
turbine engine comprising a core compressor, a recuperator, and a
turbine, the core compressor including an inlet in fluid
communication with a main compressor bleed of the main gas turbine
engine, the recuperator including a heat exchanger having a first
recuperator flow passage and a second recuperator flow passage, the
recuperator being arranged such that fluid flows in series into the
core compressor, through the recuperator first fluid flow passage,
through a combustor, through the turbine, through the recuperator
second fluid flow passage, and out an exhaust, such that exhaust
heat downstream of the turbine is transferred to combustor inlet
air; and an environmental control system inlet in fluid
communication with the main compressor bleed of the main gas
turbine engine, the environmental control system inlet comprising a
compressor drivable by power provided by the recuperated auxiliary
gas turbine engine.
Description
The present disclosure concerns an environmental control system for
an aircraft.
Prior aircraft pneumatic systems are known comprising an
environmental control system (ECS) powered by high pressure air
provided by a bleed air system (BAS). Bleed air systems generally
comprise cabin bleed offtakes which duct air from a compressor of a
gas turbine engine for use in the aircraft pneumatic system, such
as the ECS and wing de-icing. The ECS provides cabin air to the
cabin interior at a required temperature, pressure and flow rate.
The BAS system comprises a fan air heat exchanger which exchanges
heat between relatively cool fan air provided by a fan air duct,
and the relatively hot cabin bleed air provided from the engine
compressor. Once cooled by the heat exchanger, the cabin bleed air
is then passed to an air cycle machine of the ECS, where the cabin
bleed air is processed to obtain the desired heat, pressure and
flow rate, before being passed to the cabin of the aircraft. A
portion of this air is then recycled, while the remainder is vented
overboard via an outflow valve.
In some cases, low pressure and high pressure cabin bleed offtakes
are provided, which are each supplied with compressed air from the
low pressure compressor and the high pressure engine compressor
respectively, or from lower and higher pressure stages respectively
of a single engine compressor. It is generally desirable to extract
bleed air from the low pressure bleed offtake (i.e. one near the
front of the engine), since air taken from the low pressure bleed
offtake has been compressed to a lesser extent compared to air
taken from the high pressure bleed offtake. Consequently, a given
mass of air bled from the low pressure bleed offtake represents a
smaller energy loss to the thermodynamic cycle of the engine
compared to the same mass of air taken from the high pressure bleed
offtake, and so the specific fuel consumption (SFC) of the engine
will be greater (i.e. more fuel will be burned for a given thrust)
where air is bled from the high pressure bleed offtake. The bleed
air system also typically comprises one or more handling bleed
valves, which open in accordance with a schedule to control
compressor operability. Air from the handling bleeds is immediately
vented overboard without being further used.
However, current systems are non-optimal, in the sense that air
from the high pressure offtake must be used to provide sufficient
pressure and flow to the ECS at some engine conditions, such as
when the engine is at low power. Consequently, at least one of the
cabin bleed offtakes must be located at a relatively high pressure
stage of the engine compressor, and must be utilised extensively
during engine operations, particularly when the engine is operated
at low thrust, and therefore low engine overall pressure. Such
arrangements are relatively complex, and may result in excessive
thrust specific fuel consumption (SFC), since the cabin bleed air
must be provided at relatively high pressure in such prior
arrangements. In some cases, the minimum thrust that can be
achieved by the engine in flight (known as "flight idle thrust") is
limited by the overall pressure ratio required to operate the ECS
system. Consequently, in such circumstances, the engines are
operated at a higher thrust than would be required for maintaining
the desired flight profile and engine operability alone, thereby
again resulting in increased SFC.
Accordingly, it is an object of the present invention to provide an
environmental control system having improved efficiency.
According to the present invention there is provided an aircraft
pneumatic system comprising:
a main gas turbine engine comprising a main compressor;
a recuperated auxiliary gas turbine engine comprising a core
compressor having an inlet in fluid communication with a main
compressor bleed of the main gas turbine engine; and
an environmental control system inlet in fluid communication with
the main compressor bleed of the main gas turbine engine and
comprising a compressor drivable by power provided by the auxiliary
gas turbine engine.
It has been found that the above combination of features provides
an environmental control system capable of reliably providing
highly efficient pressurised air at a wide range of engine
operating conditions and aircraft operating altitudes.
The auxiliary gas turbine engine core compressor inlet may be in
direct fluid communication with the main compressor bleed of the
main gas turbine engine, or may be in fluid communication with the
main compressor bleed of the main gas turbine engine via a cabin
air exhaust. Advantageously, pressurised, waste air from the
aircraft cabin is utilised by the auxiliary gas turbine engine,
thereby increasing auxiliary gas turbine engine efficiency without
affecting main engine compressor operability.
The auxiliary gas turbine engine may comprise a variable area
turbine. Advantageously, the inlet air requirement of the auxiliary
gas turbine engine can be matched to the bleed air output of the
main engine compressor.
The main gas turbine engine compressor may comprise a high pressure
main compressor stage and a lower pressure main compressor stage.
The main compressor may comprise a high pressure compressor and a
low pressure compressor, each being driven independently by a
respective turbine. The core compressor inlet of the auxiliary gas
turbine engine may be in fluid communication with one or more of a
lower pressure main compressor customer bleed port, a higher
pressure main compressor customer bleed port, a fan outlet of the
main gas turbine engine, and a compressor handling bleed port.
In one embodiment the system may comprise a valve selectable
between a first position, in which at least one of the
environmental control system inlet and the auxiliary gas turbine
engine core compressor inlet is in fluid communication with the fan
outlet of the gas turbine engine, and a second position in which at
least one of the environmental control system inlet and the
auxiliary gas turbine engine core compressor inlet is in fluid
communication with a compressor bleed of the gas turbine engine
main compressor.
The auxiliary gas turbine engine may further comprise a load driven
by a turbine of the auxiliary gas turbine engine. The load may
comprise at least one of an electrical generator, a load
compressor, a main gas turbine engine fluid pump, a main gas
turbine engine shaft, and an air cycle machine compressor.
The environmental control system may comprise an air cycle machine
comprising, in fluid flow sequence, a compressor, a heat exchanger
and a turbine, the compressor being drivable by the turbine. The
auxiliary gas turbine load compressor may be configured to provide
air to an inlet of the air cycle machine compressor. The system may
comprise a selectable bypass arrangement configured to provide air
to the air cycle machine compressor directly from the main gas
turbine engine compressor when operated in a bypass mode, and from
the auxiliary gas turbine load compressor when in a non-bypass
mode. Where the auxiliary gas turbine engine comprises an
electrical generator, the electrical generator may be electrically
connectable to an electric motor, and the air cycle machine
compressor may be drivable by the electric motor. The air cycle
machine may comprise a first compressor driven by the motor, and a
second compressor driven by the air cycle machine turbine.
The main gas turbine engine fluid pump may comprise one or more of
a fuel pump, and oil pump and a hydraulic fluid pump.
The auxiliary gas turbine engine may be located within a main gas
turbine engine nacelle. Alternatively, the auxiliary gas turbine
engine may be located within an aircraft fuselage.
The auxiliary gas turbine engine may comprise a propelling exhaust
nozzle, which may be located within the main gas turbine engine
nacelle.
The skilled person will appreciate that except where mutually
exclusive, a feature described in relation to any one of the above
aspects of the invention may be applied mutatis mutandis to any
other aspect of the invention.
Embodiments of the invention will now be described by way of
example only, with reference to the Figures, in which:
FIG. 1 is a sectional side view of a gas turbine engine;
FIG. 2 is a schematic plan view of a first pneumatic system in
accordance with the present disclosure;
FIG. 3 is a schematic plan view of a second pneumatic system in
accordance with the present disclosure;
FIG. 4 is a schematic plan view of a third pneumatic system in
accordance with the present disclosure;
FIG. 5 is a schematic plan view of a fourth pneumatic system in
accordance with the present disclosure;
FIG. 6 is a schematic plan view of a fifth pneumatic system in
accordance with the present disclosure; and
FIG. 7 is a schematic plan view of a sixth pneumatic system in
accordance with the present disclosure.
With reference to FIG. 1, a main gas turbine engine is generally
indicated at 10, having a principal and rotational axis 9. The
engine 10 comprises, in axial flow series, an air intake 11, a
propulsive fan 12, an intermediate pressure compressor 14, a
high-pressure compressor 16, combustion equipment 18, a
high-pressure turbine 20, and intermediate pressure turbine 22, a
low-pressure turbine 24 and core and bypass exhaust nozzles 25, 27.
A nacelle 21 generally surrounds the engine 10 and defines a bypass
13 and both the intake 11 and the exhaust nozzle 20.
The gas turbine engine 10 works in the conventional manner so that
air entering the intake 11 is accelerated by the fan 12 to produce
two air flows: a first air flow into the intermediate pressure
compressor 14 and a second air flow which passes through the bypass
duct 13 to provide propulsive thrust. The intermediate pressure
compressor 14 compresses the air flow directed into it before
delivering that air to the high pressure compressor 16 where
further compression takes place.
The compressed air exhausted from the high-pressure compressor 16
is directed into the combustion equipment 18 where it is mixed with
fuel and the mixture combusted. The resultant hot combustion
products then expand through, and thereby drive the high,
intermediate and low-pressure turbines 20, 22, 24 before being
exhausted through the core nozzle 25 to provide additional
propulsive thrust. The high 20, intermediate 22 and low 24 pressure
turbines drive respectively the high pressure compressor 16,
intermediate pressure compressor 14 and fan 12, each by suitable
interconnecting shaft 30, 28, 26. The high pressure shaft 26
further drives a main gas turbine engine electrical generator 57.
Consequently, the main gas turbine 10 provides primary motive power
for an aircraft 1.
FIG. 2 shows a first pneumatic system of the aircraft 1 in
accordance with the present disclosure. The system 100 includes a
gas turbine engine 10 similar to that of FIG. 1. The gas turbine
engine comprises a first bleed offtake 32 in fluid communication
with a compressor stage of the intermediate pressure compressor 14.
The offtake supplies compressed air to both a handling bleed valve
34 and a customer bleed air valve 36 via a common manifold. The
handling bleed valve 34 is operable in either fully open and fully
closed positions, and is operated by a compressor handling schedule
governed by an engine controller (not shown). An exhaust of the
handling bleed valve 34 vents to the engine nacelle 21.
The customer bleed air valve 36 is operable between fully closed
and fully open, as well as intermediate positions, and provides air
downstream to a plurality of components. A first component supplied
by the customer bleed air valve 36 is a recuperated auxiliary gas
turbine engine generally indicated at 40. The auxiliary gas turbine
engine comprises, in fluid flow sequence, a core compressor 42, a
recuperator in the form of a heat exchanger 44, a combustor 46 and
a turbine 48. The compressor 42 and turbine 48 are interconnected
by a shaft 50, such that rotation of the turbine 48 drives the
compressor 42, and the auxiliary gas turbine 40 consequently
operates in a similar manner to the main gas turbine engine 10. The
auxiliary gas turbine engine 40 is provided within the main engine
10 nacelle.
The turbine 48 is optionally a variable area turbine comprising
variable inlet guide vanes (indicated by the arrow), which are
adjustable to vary the inlet area of the turbine 48. Consequently,
auxiliary gas turbine engine power output can be varied while
maintaining a substantially fixed overall pressure ratio, by
varying fuel flow and turbine 48 inlet area. Alternatively, airflow
rate can be varied independently of engine 40 power output by
varying the inlet area of the turbine 48. Operation in this manner
is greatly advantageous in the present invention, for reasons
outlined below.
The auxiliary gas turbine engine 40 is recuperated, as explained in
detail below. The heat exchanger 44 comprise a first flow passage
52 in direct fluid flow communication with the core compressor 42
and combustor 46, and a second flow passage 54 which is in direct
fluid flow communication with an outlet of the turbine 48 and an
exhaust outlet. Consequently, fluid flowing through the auxiliary
gas turbine 40 flows in series into the compressor 42, through the
recuperator first fluid flow passage 52, through the combustor 46,
through the turbine 54, through the recuperator second fluid flow
passage 54, and out the exhaust. Consequently, waste heat from the
auxiliary gas turbine exhaust is recycled prior to combustion,
thereby increasing efficiency. In view of this arrangement, the
fuel flow of the auxiliary gas turbine engine 40 is reduced for a
given power output compared to a simple cycle gas turbine engine.
Meanwhile, the auxiliary gas turbine engine 40 can operate at high
efficiency with various constraints such as variable inlet
temperature, pressure and flow rates as well as variable loads in
view of the recuperation and/or variable area turbine. Airflow from
the turbine 48 is exhausted to either the engine core nozzle 25 or
the fan nozzle 27.
The auxiliary gas turbine engine 40 comprises at least one load. In
this case, an electrical generator 56 and a load compressor 58 are
provided. The electrical generator 56 comprises an alternating
current (AC) electrical generator which is electrically coupled to
aircraft and engine electrical loads such as fuel pumps, hydraulic
fluid pumps, aircraft auxiliary systems, etc.
The load compressor 58 comprises an air compressor 58. The
compressor 58 is provided with air from the customer bleed 32 via
the customer bleed air valve 36, and is driven by the turbine 48
via the shaft 50. Consequently, the load compressor is driven by
power provided by the auxiliary gas turbine engine 40. In
alternative embodiments, the load compressor 58 could be driven by
a separate auxiliary gas turbine engine turbine, which may be
interconnected by a separate shaft. Such an arrangement is known as
a "free power turbine", as the separate turbine does not drive a
compressor of the engine core. An outlet of the air compressor 58
is provided to an inlet of an Environmental Control System (ECS)
79, as described in further detail below.
The system further comprises a manifold 70 having a first inlet
provided with low pressure fan air via a first non-return valve 72,
and a second inlet provided with high pressure from the main gas
turbine engine intermediate pressure compressor 14 via the customer
bleed air valve 36. The manifold 70 has a first output which
provides air from either the first or second input to inlets of the
auxiliary gas turbine engine core compressor 42 and load air
compressor 58. A second output of the manifold 70 provides air from
either the first or second input to the inlet of the ECS 79 via an
air passage 62. The system is operable between a first mode, in
which air is provided from the fan 12 outlet to the compressors 42,
58 inlets, a second mode, in which air is provided from the
intermediate pressure compressor 14 to the compressors 42, 58 and a
third mode, in which air is provided from the intermediate
compressor 14 directly to the ECS inlet. Operation in accordance
with the first, second and third modes is described in further
detail below.
The ECS 79 comprises a pre-cooler heat exchanger 60. The pre-cooler
60 comprises a first flow passage 62 which is supplied with
compressed air from either the load compressor 58 via a non-return
valve 64 when the system is operating in either the first or second
mode, or directly from the intermediate pressure compressor 14 via
a non-return valve 74 and the customer bleed air valve 36 when the
system is operated in the third mode. (when the selector valve 70
is in the third position). The pre-cooler 60 comprises a second
flow passage 66 through which lower pressure, cooler air is
provided from the fan 12 outlet via a flow control valve 68. The
pre-cooler and flow control valve 68 allows the temperature of air
flowing into the ECS 79 to be controlled.
The passage 62 leads to a first ECS heat exchanger 76 provided
downstream, which exchanges excess between warm air within the
passage 62 and cooler air provided from a ram air inlet 78 provided
on an external surface of the aircraft fuselage or nacelle. Cooled
air within the passage 62 is provided to an air cycle machine 75
downstream of the further heat exchanger 76. Air is directed
through the air cycle machine 75, in fluid flow series, through an
air cycle machine compressor 80, a second ECS heat exchanger in the
form of an air cycle machine heat exchanger 82, a third heat
exchanger 88, a fourth heat exchanger 90, a water extractor 92,
back through the third heat exchanger 88, through an air cycle
machine turbine 84, back through the fourth heat exchanger 90, and
on to a cabin 94. Consequently, the air is first pressurised and
thereby heated by the compressor 80. This heated air is cooled by
the second heat exchanger 82 by exchanging heat in the compressed
air with relatively cool air from the ram air inlet 78. This cooled
air is then passed to the third and fourth heat exchangers, which
cool the air such that humidity within the air condenses to be
removed by the water extractor. The cooled air is then re-heated by
passing through the fourth heat exchanger once more, before being
cooled again by passing through the turbine 84, thereby driving the
turbine 84, which in turn drives the compressor 80 via an
interconnecting shaft 81. This cool air is used to warm air passing
through the fourth heat exchanger on its first pass, thereby being
warmed again before passing to the cabin 94.
As explained above, air is cooled by the air cycle machine 75, and
passed to the aircraft cabin 94. A portion of the air from the
cabin 94 is recycled through the air cycle machine 75, while the
remainder is exhausted from the aircraft through an outflow valve
96, providing some thrust.
The system is operated as follows. The system may be operated in
accordance with one of a first, second or third operating mode,
depending on engine and aircraft operating conditions.
The system is generally operated in the first operating mode when
the main gas turbine engine 10 is inoperative, such as during
embarkation when on the ground. In the first operating mode, the
bleed air valve is shut, such that air is provided only from the
fan 12 outlet to the inlets of the auxiliary gas turbine engine
core compressor 42 and the load compressor 58. Fuel is provided to
the auxiliary gas turbine engine combustor 46, such that the
auxiliary gas turbine engine 40 provides motive power, to thereby
operate the load compressor 58. Compressed air from the load
compressor 58 flows through the first flow passage 62 of the
pre-cooler 60. The flow control valve 68 may be shut, thereby
preventing flow from the fan 12 flowing through the pre-cooler 60,
such that air entering the ECS 79 is not significantly cooled by
the pre-cooler 60. One way valve 74 prevents fluid flow from
re-entering the manifold 70 after being compressed by the load
compressor 58, thus preventing recirculation. This air is then
passed to the air-cycle machine, which operates as described above
to regulate cabin air temperature and pressure. Meanwhile, the
auxiliary gas turbine engine 40 also provides motive power to the
electrical generator 56, thereby providing electrical power for the
aircraft 1.
Once the main gas turbine engine 10 is started, the system may be
operated in a second operating mode. The system is typically
operated in the second operating mode when the intermediate
pressure compressor 14 is operated at a low pressure ratio, such as
during descent and taxiing. In the second operating mode, customer
bleed air valve 36 is opened such that air is provided from
intermediate pressure compressor 14 via the customer bleed air
valve 36 to the inlets of the auxiliary gas turbine engine core
compressor 42 and the load compressor 58. One way valve 72 prevents
this flow from being lost to the fan inlet 72. Fuel is again
provided to the auxiliary gas turbine engine combustor 46, such
that the auxiliary gas turbine engine 40 provides motive power, to
thereby operate the load compressor 58 and generator 56. Again,
one-way valve 74 prevents flow recirculation. Compressed air from
the load compressor 58 flows through the first flow passage 62 of
the pre-cooler 60, and cooling air is provided through the second
flow passage 66 from the nacelle 13. Fan air flow rate is
controlled by valve 68 to provide a desired pre-cooler outlet
temperature. This air is then passed to the air-cycle machine,
which operates as described above to regulate cabin air temperature
and pressure.
In the third operating mode, the bleed air valve 36 is opened, and
air is provided from intermediate pressure compressor 14 via the
customer bleed air valve 36 directly to the ECS pre-cooler 60, i.e.
without being compressed by the load compressor 58. In this case,
air is optionally provided to the auxiliary gas turbine engine 40,
and so the load compressor 58 and the generator 56 are optionally
operated. If the load compressor 58 is not operated, non-return
valve 64 will operated to prevent air flow running backwards
through the load compressor to the manifold 70, and thereby
recirculating. Again, air is fed through the pre-cooler from the
fan 12 outlet, with the flow rate being controlled by valve 68 to
provide a desired pre-cooler outlet temperature. In this scenario,
the system operates similarly to a conventional bleed air
environmental control system, with air pressure being provided
solely by the main gas turbine engine 10, and electrical power
being provided by a main gas turbine engine electrical generator
(not shown).
Consequently, the system can be operated in one of a first, second
or third mode by varying the position of a single valve 36, and
either operating or not operating the auxiliary gas turbine engine
load compressor 58. Consequently, only a single variable valve is
provided for switching between operating modes, which is one fewer
than current systems employing high pressure and low pressure
compressor bleeds and corresponding valves. Consequently, the
system has reduced cost and complexity, and increased
reliability.
Consequently, the system is highly flexible, providing three
operating modes suitable for providing high efficiency ECS input
air at different phases of flight. In general, the system is as
efficient or more efficient than previous systems, while having the
same or lower weight in view of the closer integration between the
auxiliary and main gas turbine engines, and the deletion of the
high pressure customer bleed.
A further benefit of the present invention is that, due to the
provision of an inlet to the auxiliary gas turbine engine from the
main gas turbine engine compressor bleed, a the auxiliary gas
turbine engine can be operated under circumstances in which a
conventional APU could not be (for example, at high altitude).
Consequently, an increased safety factor can be provided, or one of
the redundant electrical generators (normally at least two per
engine are provided) could be deleted, thereby saving weight.
FIG. 3 shows a second embodiment of the present disclosure. Similar
parts are given corresponding reference numerals, incremented by
100. Again, a main gas turbine engine 110 is provided, having a fan
112, intermediate and high pressure compressors 114, 116, combustor
118, high, intermediate and low pressure turbines 120, 122, 124,
and suitable interconnecting shafts. A recuperated auxiliary gas
turbine 140 is provided, which is again similar to the auxiliary
gas turbine engine 40, having a core compressor 142, heat exchanger
144, combustor 146 and turbine 154 interconnected by a shaft 150.
However, in this case, no load compressor is coupled to the shaft
150, with only an electrical generator 156 being provided as a load
to the auxiliary gas turbine engine 140.
The system further comprises an environmental control system
pre-compressor 158, which performs a similar function to the load
compressor 58, but is driven by an electrical motor 198 instead of
the auxiliary gas turbine engine 140. The pre-compressor 158 is
supplied with air from the intermediate pressure compressor 114 via
a customer bleed valve 136, and provides compressed air to the
pre-cooler 160. Again, air from the pre-cooler is supplied to the
environmental control system downstream of the pre-cooler 160.
The system further comprises a manifold, through which air flows in
accordance with first, second or third modes. In operation, when
the system is operated in one of the first or second mode, air is
supplied to both the auxiliary gas turbine engine 140 and the
pre-compressor 158 (from the fan 112 in the first position, and
from the intermediate pressure compressor 114 in the second
position). The first mode is selected by closing the bleed valve
136, and the second mode is selected by opening the bleed valve
136. The pre-compressor 158 is driven by the electrical motor 198
to pre-compress air to a desired level to provide air to the ECS
79. In the third mode, the bleed valve 136 is opened, and the
pre-compressor 158 is not operated. Consequently, air is provided
to the ECS 79 directly from the main engine compressor 114. The
electrical motor 198 can be operated either by electrical power
provided by the generator 156 driven by the auxiliary gas turbine
engine 156, or the electrical generator 157 driven by the main gas
turbine engine 10. Consequently, this embodiment is more flexible
than the first embodiment, since the compressor 158 is not
dependent on operation of the auxiliary gas turbine engine.
Furthermore, since there is no physical shaft between the auxiliary
gas turbine 40 and the pre-compressor 158, these components do not
have to be co-located, though the pre-compressor is preferably
located within the main gas turbine engine nacelle.
The electric motor comprises a variable speed electric motor 198.
Consequently, the pressure provided to the ECS 79 can be varied
readily and accurately.
FIG. 4 shows a third embodiment of the present disclosure having
similar main 210 and auxiliary 240 gas turbine engines. This
embodiment is similar to the second embodiment, having a
pre-compressor 258 driven by an electric motor 298. However, in
this case, the pre-compressor 258 is provided downstream of the
pre-cooler 260, and is located within a fuselage 2 of the aircraft
1. Again, the pre-compressor 258 provides compressed air to
downstream components of the ECS, in particular the air cycle
machine. Again therefore, ECS inlet pressure can be modulated to an
appropriate level by the pre-compressor 258 using electrical power
from either the auxiliary gas turbine engine driven generator 256
or the main gas turbine engine generator 257. As shown by the
dashed lines in FIG. 4, the electrical generator 256 is
electrically connectable to the electric motor 298, which drives
air cycle machine compressor 258.
Again, the system can be operated in one of a first, second and
third operating mode. In a first operating mode, air is provided
from the fan 212 to the ECS via a manifold 270 by closing a main
engine compressor bleed valve 236 an operating the auxiliary gas
turbine engine combustor 246. In this mode, air is also provided
from the fan 212 to the auxiliary gas turbine engine core
compressor 242 to provide air to the auxiliary gas turbine engine
240. In a second mode, selected by opening the bleed valve 236 and
operating the combustor 246, air is provided from the intermediate
compressor 214 to both the ECS and the auxiliary gas turbine engine
core compressor 242. In the third mode, selected by opening the
valve 236 and shutting down the combustor 246, air is provided from
the intermediate pressure compressor 214 to the ECS only.
FIG. 5 shows a fourth embodiment of the present disclosure, which
is similar to the third embodiment, however in the fourth
embodiment, the pre-compressor is omitted entirely. Instead, the
air cycle machine compressor 380 is coupled to an electrical motor
398, which is in turn selectively provided with power from either
the auxiliary gas turbine engine driven generator 356 or the main
gas turbine engine driven generator 357.
The fourth embodiment is operated in accordance with first, second
or third modes similar to those of the third embodiment, but with
the air cycle machine compressor 380 being driven by the electrical
motor 398 in place of the pre-compressor. Such an arrangement
requires fewer modifications compared to previous versions, though
the air cycle machine compressor may require modification.
FIG. 6 shows a fifth embodiment of the disclosure. In this
embodiment, main and auxiliary gas turbines 410, 440 are provided,
which are again similar to those in previous embodiments. A
pre-compressor 458 is provided, which is located within the main
gas turbine engine nacelle, and is selectively connected to the fan
412 and intermediate pressure compressor 412, 414 of the main gas
turbine engine 410 and the ECS via a manifold 470. Again, the
pre-compressor 458 can be supplied with air either from the fan 412
or main engine compressor 414.
The auxiliary gas turbine engine 440 is located within the fuselage
2 in this embodiment. An inlet of the auxiliary gas turbine engine
core compressor 442 is provided within air from the cabin 494 after
it has been exhausted, prior to be vented from the aircraft via the
outflow valve 496. Consequently, air is supplied to the auxiliary
gas turbine core compressor 442 indirectly from either the gas
turbine engine fan 412 or the intermediate pressure compressor 414
via the cabin 494. Consequently, the pressurised air from the cabin
494 is utilised by the auxiliary gas turbine engine 440, thereby
further raising the efficiency of the system as a whole. Since the
pre-compressor 458 is in this case electrically driven by a motor
498, the auxiliary gas turbine engine 440 and pre-compressor 458
again do not have to be co-located.
FIG. 7 shows a sixth embodiment of the disclosure. The system shown
in FIG. 7 is similar to that of FIG. 2, having a main engine 510
and auxiliary engine 540. However, the auxiliary engine is utilised
to drive additional equipment in addition to the load compressor
558 and generator 556. The additional equipment includes main
engine 10 fluid pumps including a main engine hydraulic pump 599a,
fuel pump 599b and oil pump 599c. Each of the pumps 599a-c and the
generator are driven by the auxiliary gas turbine engine shaft 550,
via a gearbox 597. Consequently, the high efficiency of the
auxiliary gas turbine engine 540 is utilised for a wider range of
uses, thereby further reducing overall system fuel burn.
It will be understood that the invention is not limited to the
embodiments above-described and various modifications and
improvements can be made without departing from the concepts
described herein.
For example, it will be appreciated that the disclosure is equally
applicable to two shaft gas turbine engines, employing a fan and a
low pressure compressor mounted to a common shaft, and a high
pressure compressor mounted to a further shaft. The auxiliary gas
turbine engine may mechanically drive the air cycle machine
compressor. The electric generator and motor could be DC. The
electric motor could be constant speed, with the speed of the
driven load being varied by a variable speed gearbox such as a CVT,
or the pressure or flow rate of the compressor could be varied by a
bypass arrangement. The system could comprise a second, high
pressure customer bleed for redundancy or further increased
flexibility.
Except where mutually exclusive, any of the features may be
employed separately or in combination with any other features and
the disclosure extends to and includes all combinations and
sub-combinations of one or more features described herein.
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